This disclosure relates to a gas turbine engine component, such as a turbine airfoil. Particularly, the disclosure relates to a film cooling hole used to communicate fluid from an internal passageway to an exterior surface.
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
The advancement of turbomachinery performance is linked to both the over-all pressure ratio of the machine and the turbine inlet temperature that can be reliably sustained during service. Increases in efficiency through either of these methods typically produces a hotter operating environment for turbine flow path hardware in which the working fluid is typically several hundreds of degrees hotter than the melting point of the component alloys. Dedicated cooling air is extracted from the compressor and used to cool the gas path components in the turbine, which can incur significant cycle penalties.
For extremely high temperature applications, film cooling is typically utilized along with backside convection. This method uses cooling air delivered internal of the component and expelled through holes in the exterior airfoil surface to provide a cooling flow over the external surface that reduces the local external surface temperatures downstream. Typically cooling holes are machined into the part and are round or diffuser shaped as permitted by a typical laser or EDM machining process.